In axial flow compressors, as employed in conventional gas turbine engines, the Mach number of the flow relative to the individual rotor and/or stator blades is typically in the subsonic or transonic flow regime. Blade tip Mach numbers from about 0.5 Mach to about 0.7 Mach are common. Rotor/stator operation at Mach numbers in this range results in high lift to drag levels, with minimal shock losses for the rotor and stator blades. However, one of the disadvantages of such a design is that the pressure ratio that can be achieved across any given rotor/stator stage is typically limited to about 1.5:1. Yet, simple gas turbine systems that must achieve high cycle efficiency levels require overall compression ratios in excess of about 10:1. Further, systems with compression ratios up to about 25:1 have been demonstrated for applications with demanding performance requirements. Consequently, when there is a demand for high compression ratios, compressors with many stages of compression are provided. Unfortunately, such multi-staged compressors are relatively heavy, complex, and thus are expensive. As a result, there has been a continuing interest in the compressor design field to explore higher rotor/stator loadings, in order to deliver high compression ratios with fewer stages.
Further, the limitations imposed by low stage pressure ratios in subsonic compressors have stimulated the study, design, development, and testing of transonic and supersonic flow velocities in the rotor blades. Such a design approach provides a much greater amount of kinetic energy to the gas at each stage. With supersonic compressors of axial or mixed flow (i.e. using combinations of axial and radial flow), past designs have shown the potential to attain stage pressure ratios as high as 4 or even 6, with attendant adiabatic efficiencies of 75% to 80%. In such designs, it follows that high air handling ability is obtained with minimal frontal area, which is particularly important for flight applications. Furthermore, high pressure ratios per stage means that fewer stages are required, with resultant saving in compressor weight and expense.
High compression ratios per stage have been provided by several types of designs. In supersonic designs, shock waves may be handled in the stator, or in the rotor, or both. No matter where the shocks occur, the basic design criteria is to minimize the pressure loss and to insure flow stability over as wide of an operating range as possible. One type of prior art blade configuration that accomplishes shock compression within the rotor flow is illustrated in FIG. 1. There, a rotor operating at an inlet Mach number of 1.8 provides a first rotor blade R1 which generates a first shock S1 which is captured and reflected in the form of a second, oblique shock S2 by the adjacent rotor blade R2. Resultant downstream flow decelerates behind the third shock S3 to a Mach number of 0.75, and, after expansion to a Mach number of 0.5. For achievable supersonic blade tip Mach numbers, considerable static pressure rise is obtained in the rotor itself. Yet, one of the disadvantages of such a rotor blade design has been the loss incurred in the subsonic diffusion process, due to separation occurring from interaction between the normal shock S3 and the boundary layer.
As an initial step in an attempt to overcome the shortcomings inherent in presently available compressor designs, we have evaluated the performance of various supersonic flight inlets. Most manned aircraft and many missiles rely on some form of air-breathing propulsion for sustained flight within the earth's atmosphere. Air breathing engines require an inlet to diffuse air from the free-stream velocity to a lower velocity that is acceptable for further processing by other engine components. The inlet components are designed to capture the exact amount of air required, and to accomplish diffusion with a minimum of total pressure loss. Importantly, such inlets also must deliver the air to the subsequent components within acceptable levels of flow distortion. Such inlets must also be configured to contribute the lowest amount of external drag possible for a given application. Because a wide range of supersonic air-breathing propulsion systems have been designed, developed, and tested over the last 60 years, the optimization of supersonic inlet configurations for various flight applications has received a great deal of attention. As a result, many techniques are known in that art for maximizing the performance of supersonic inlets. In fact, performance levels of such inlets have been well established over a wide range of operational flight Mach numbers.
Attention is directed to FIG. 2, wherein the supersonic inlet performance (as represented by the total pressure recovery) is shown as a function of flight Mach number, for a wide range of supersonic inlet systems. The various inlet performance data points shown on this plot represent actual test data from supersonic inlet systems of many different configurations that have been designed to operate over a wide flight Mach number range, while exposed to a range of angle of attack attitudes and yaw angles. Examples are provided for specific designs by Pratt & Whitney (P&W), United Technologies (UTRC), the NASA Hypersonic Research Engine (HRE), the US Army Transportation Material Command (TMC), and the National Aeronautics and Space Administration (NASA) TMX 413 design. Additionally, the nominal inlet performance requirement set forth in United States Military Specification MIL-E-5007 is illustrated. In short, the peak performance levels of each of the noted systems have been compromised to achieve the robust operability and stability requirements of a true flight system that must complete a mission wherein a wide range of inlet flow conditions are encountered.
For categorizing anticipated performance characteristics, and for evaluation and comparison purposes, supersonic flight inlet designs are often defined into three broad groups. These groups are (a) normal shock inlets, (b) external compression inlets, and (c) mixed compression inlets. These three different groups are depicted in FIG. 3, along with a relative representation of performance levels for each group as a function of design Mach number. As can be appreciated by reference to FIG. 3, each of these three types of inlets has advantages and disadvantages. The normal shock inlet exhibits excellent pressure recovery at relatively low Mach number, but recovery drops markedly as design Mach number increases. The external compression inlet shows good pressure recovery in the Mach 2 range, but also drops off markedly as the operating Mach number increases above this design range. The mixed compression inlet provides acceptable pressure recovery over a relatively broad range of design Mach numbers, but again, efficiency drops off markedly as the operating Mach number reaches 4 or more.
Referring again to FIG. 2, performance data is included for all three types of inlet that were depicted in FIG. 3, as can generally be appreciated by the range provided for flight Mach numbers across which the various designs operate. Importantly, it can be appreciated that a properly designed supersonic inlet that is optimized for operation at a single Mach number (or a small operating Mach number range), with minimal variability in angle or attack or in yaw exposure, could achieve performance levels somewhat greater than the performance levels indicated in FIG. 2.
Further, in FIG. 4, nominal inlet performance requirements as delineated in Military Specification MIL-E-5007 are provided. First, this figure provides a curve that corresponds to the mean line of the flight inlet performance, expressed as static pressure ratio, derived from the curve for the MIL-E-5007 inlet depicted in FIG. 2. Second, based on the first derived curve, FIG. 4 provides a curve that corresponds to the mean line of flight inlet performance expressed as adiabatic (isentropic) compression efficiency as a function of flight Mach number. For example, at a flight Mach number of 2.4, the MIL-E-5007 specification inlet will provide compressor performance at 94% adiabatic efficiency.
Total pressure recovery is commonly used by flight inlet designers to evaluate the performance of such systems. The total pressure recovery is the ratio of the total pressure of the flow leaving the inlet to the free stream total pressure level. The flight Mach number can also be thought of as quantifying the magnitude of the total pressure available to an inlet. Thus, it is possible for any given flight Mach number and total pressure recovery level to calculate a corresponding system pressure ratio. The pressure ratio is a critical factor for all compression applications.
In FIG. 5, the compression efficiency for the flight inlet data shown in FIGS. 2 and 4 has been averaged, and reduced to a single functional line where adiabatic compression efficiency is illustrated as a function of inlet static pressure ratio. Thus, FIG. 5 represents the basic efficiency characteristics of a wide range of flight inlets as a function of static pressure ratio. Importantly, this comparison allows the generalized efficiency of flight inlets to be directly compared to the performance of (a) centrifugal compressors, and (b) axial flow compressors, in terms of adiabatic (isentropic) compression efficiency. To facilitate this comparison, the adiabatic (isentropic) compression efficiency of a number of selected axial flow industrial gas turbine compressors are shown in this FIG. 5. It can readily be appreciated from FIG. 5 that supersonic flight inlet designs have the potential to operate at significantly greater efficiencies than heretofore known centrifugal compressors or conventional axial flow turbo-compressors.
The isentropic compressor efficiency (sometimes referred to as the adiabatic compressor efficiency) is a performance parameter commonly used by compressor designers. This is based on a theoretical frictionless adiabatic compression process, and thus also is known as isentropic, or having a constant entropy. Although it is evident that compression is not frictionless, and that friction results in heating of the metal parts of the compressor, the assumption that adiabatic compression takes place may be used in computing the theoretical power requirements for a particular compression requirement. The isentropic compressor efficiency is defined as the ratio of isentropic work of compression to the actual work of compression. Equation 1 shows the definition for the isentropic compressor efficiency:ηcomp=(ht2i−ht1)/(ht2a−ht1)
Although a variety of supersonic gas compressor and compressor diffuser designs have been heretofore proposed, in so far as we are aware, none have been widely utilized for primary compressor service, whether for common applications such as air or steam, or heavier gases such as certain refrigerants, or heavy chemical intermediates such as uranium hexafluoride. Undoubtedly, improvements in compression efficiency, which would be especially advantageous in order to reduce energy costs for a particular compression service, would be desirable. In various attempts to achieve such improvements, many different methods and structures have been tried, either experimentally or commercially. Some of such attempts have included the use of various shock patterns, such as adapted from conventional centrifugal compressor wheels, or incorporating various multiple rotor configurations. The challenge, however, has been in the selection of methods and structures that assure adequate performance (including such acceptability of attributes such as starting performance, overall efficiency, and acoustic stability) while reducing capital and operating costs. It would be especially desirable for compressors operating under such conditions to have an inlet and diffuser configuration that would be resistant to small changes about a design point with respect to external flow field dynamics and shock perturbations.
Consequently, it would be desirable to provide a reliable supersonic gas compressor, specifically including a compressor and diffuser chamber structure that enables the compressor to maintain high isentropic efficiency with a minimum of structure, while operating at a high pressure ratio. Therefore, a continuing demand exists for simple, highly efficient and inexpensive gas compressors as may be useful in a wide variety of gas compression applications. This is because many gas compression applications could substantially benefit from incorporating a compressor that offers a significant efficiency improvement over currently utilized designs. In view of ever increasing energy costs, particularly for both electricity and for natural gas, it would be desirable to attain significant cost reduction in utility expense for gas compression. Importantly, it would be quite advantageous to provide a novel compressor which provides improvements (1) with respect to operating energy costs, (2) with respect to reduced first cost for the equipment, and (3) with respect to reduced maintenance costs. Fundamentally, particularly from the point of view of reducing long term energy costs, this would be most effectively accomplished by attaining gas compression at a higher overall compression efficiency than is currently known or practiced industrially. Thus, the important advantages of a new gas compressor design providing the desirable feature of improved efficiency can be readily appreciated.